As is known, an axial compressor for a gas turbine engine may include a number of stages arranged along an axis of the compressor. Each stage may include a rotor disk and a number of replaceable compressor blades arranged about a circumference of the rotor disk. To facilitate replacement, the blades may be removably attached to the rotor disk via dovetail connections by which root portions of the blades are inserted axially into respective slots formed about the circumference of the rotor disk. According to a full-pitch platform configuration, each blade may include a platform portion extending circumferentially and abutting the platform portions of adjacent blades. In this manner, the platform portions may define a radially inner boundary of a compressed air flowpath. Additionally, the platform portions may define a radially outer boundary of a cavity formed between the platform portions and an outer surface of the rotor disk. During operation of the compressor, a portion of the compressed air may pass upstream through the cavity from a high-pressure side of the compressor blades to a low-pressure side of the compressor blades. Such stage-to-stage leakage of compressed air may reduce efficiency and surge margin of the compressor itself as well as the overall gas turbine engine.
Certain axial compressors including compressor blades having a full-pitch platform configuration may include a cover plate positioned over the cavity on at least one of the upstream side or the downstream side of the blades. In this manner, the cover plate may reduce stage-to-stage leakage of compressed air, although the cover plate and associated hardware may increase the complexity, size, and weight of the compressor stage at the disk-blade interface. Other axial compressors may reduce stage-to-stage leakage by including a sealant, such as a room temperature vulcanizing (RTV) sealant, which fills at least a portion of the cavity to block air flow therethrough. However, such a sealant may be difficult to design and validate for long-term leakage control in an axial compressor because it may degrade over time and thus may allow for varying levels of leakage over the life of the compressor.
There is thus a desire for an improved axial compressor for a gas turbine engine and a method for controlling stage-to-stage leakage therein. Specifically, such a compressor may control leakage of compressed air through a cavity formed between a rotor disk and platform portions of compressor blades having a full-pitch platform configuration. Such leakage control may increase efficiency and surge margin of the compressor and the overall gas turbine engine. Preferably, such a compressor will not require additional components at the disk-blade interface or a sealant that may degrade over time.